Environmentally stable hybrid fabric system for exterior protection of an aircraft

ABSTRACT

A method of forming an exterior surface protective structure ( 12 ) for an aircraft ( 10 ) includes uniting a loaded surfacer ( 52 ) having a carrier ( 51 ) to a hybrid prepreg substrate ( 32 ). The prepreg substrate ( 32 ) includes a carbon fabric ( 44 ) with an integral conductive component ( 48 ) having conductivity with in a metal conductivity range and is united to a base substrate ( 30 ). The loaded surfacer ( 52 ) and the prepreg substrate ( 32 ) are cured, which includes interfacially adhering matter between the loaded surfacer ( 52 ) and the prepreg substrate ( 32 ). A protective fabric system ( 12 ) for an exterior ( 14 ) of an aircraft ( 10 ) includes the base substrate ( 30 ). The hybrid prepreg substrate ( 32 ) is coupled to the base substrate ( 30 ). The loaded surfacer ( 52 ) with the carrier ( 51 ) is interfacially adhered to the prepreg substrate ( 32 ).

RELATED APPLICATION

This application is a divisional application of, and claims priorityfrom prior application Ser. No. 11/163,614, filed Oct. 25, 2005.

TECHNICAL FIELD

The present invention is related generally to aircraft exteriorcoatings, layers, surfaces and composites. More particularly, thepresent invention is related to a system for exterior protection of anaircraft that provides corrosion resistance, rain erosion resistance,environmental durability, structural performance, and electromagneticprotection including lightning protection.

BACKGROUND OF THE INVENTION

Traditionally, to protect against lightning aircraft methods haveincluded a low resistance pathway throughout the metallic bulk of thefuselage to dissipate the electrical energy. Metallized fiber reinforcedstructural materials have been used along the exterior surfaces ofcomposite parts to provide a medium to rapidly dissipate the energy.Some of the present lightning protective structures, although feasiblefor use on spacecraft and some aircraft, are not feasible for use onhigh use commercial aircraft. This is due to the rigorous andcontinuously changing pressure, humidity, and temperature environmentexperienced by commercial aircraft, as well as the different cost andmaintenance constraints associated therewith.

Testing has shown that under high use commercial aircraft operatingconditions certain lightning protective structures tend to experiencesubstrate microcracking and finish cracking making them more susceptibleto corrosion and ultraviolet degradation. Microcracking is sometimesreferred to as “weave telegraphing.” Weave telegraphing refers to when:(a) the visual irregularities in the finishes take on the appearance ofthe underlying weave pattern of the surface, (b) the pattern becomesmore pronounced while in-service, and (c) there is formation andpropagation of substrate and/or paint finish cracking. The statedmicrocracks tend to form due to repeated and extreme temperature,humidity, and pressure fluctuations. Microcracking occurs due to anumber of factors including internal stresses from differences incoefficient of thermal expansion, as well as from non-optimum interfaceadhesion between components in composite systems.

The microcracks can extend into visual paint layers, which can result inappearance degradation and increased maintenance and inspection timesand costs. increased maintenance such as paint repair, is needed notjust for appearances, but also to identify when repainting is necessaryto prevent ultraviolet degradation of the underlying organic materials.Increased inspection is needed not only to monitor corrosion, but alsoto ensure the microcracks have not adversely affected structuralintegrity. Thus, such structures are not always cost-effective forlong-term use in the commercial environment.

One type of lightning protective structure includes a substrate layer, ametal mesh screen, and a non-structural outer film that may bereinforced with materials such as glass or polyester. The mesh can be ametal woven fabric, random mat, or perforated metal that is usuallyexpanded. Depending on the metal and substrate an additionalnon-structural prepreg layer may be used for galvanic isolation to avoidcorrosion between the base substrate and the metal mesh. Although thisstructure provides the desired lightning protection, it provides nostructural benefit and contains multiple non-structural layerstypically. Thus, the structure is inherently labor intensive and costlyto produce. The weight of resin needed to encapsulate the mesh toprevent corrosion and provide a smooth surface can exceed the weight ofthe metal mesh and as a result is heavy. Also the mesh system can besusceptible to microcracking.

Another protective structure approach is to use a solid metal overcomposite material. This structure is also heavy and difficult toprocess without manufacturing defects, such as voids, when co-cured as asolid film or applied as a spray to the cured part. Spray processes suchas aluminum flame spray have the added complication of requiringqualified personnel and equipment typically not available at airlinefacilities.

Thus, there exists a need for an improved lightning protective structurefor an aircraft that does not exhibit the above-mentioned disadvantagesand provides the corrosion resistance, rain erosion resistance,environmental durability, structural performance, and electromagneticprotection including lightning protection characteristics desired.

SUMMARY OF THE INVENTION

One embodiment of the present invention provides a method of forming anexterior surface protective layer for an aircraft by uniting a loadedsurfacer having a carrier to a hybrid prepreg substrate. The hybridprepreg substrate includes carbon fabric with an integral conductivecomponent that is united to a base substrate. The loaded surfacer andthe hybrid prepreg substrate are cured at the same time, which includesinterfacially adhering matter between the loaded surfacer and the hybridprepreg substrate. When co-cured with the base substrate processingcosts are reduced.

The hybrid substrate of the above-stated embodiment provides increasedstructural durability, as well as electromagnetic protection. There areno supplemental electromagnetic protection costs associated with thehybrid substrate because the protection is integral to the structuralhybrid prepreg thereof.

The embodiments of the present invention provide several advantages. Onesuch advantage is the provision of using an inorganic filler loadedsurfacer over a hybrid fabric substrate, which are commingled duringcure. The hybrid fabric substrate provides the desired lightningprotection while the combination of the loaded surfacer and the hybridfabric substrate provide the desired structural support andenvironmental protection. The use of the loaded surfacer inhibits bothcorrosion and microcracking and thus improves durability.

Another advantage provided by an embodiment of the present invention, isthe provision of evenly distributing a loaded surfacer over the surfaceof a hybrid prepreg substrate and interfacially adhering that surfacerto that substrate. The surfacer which has a carrier and fillerscompensates for pressure differences due to design or tooling featuresthat otherwise would promote mark-off and lead to large variations insurfacer thickness. This assists in preventing the type of weavetelegraphing that leads eventually to substrate microcracks and paintcracks. The above-stated advantages in combination provide a protectivestructure that is capable of withstanding the qualification testing andoperating environment for high use commercial aircraft.

The present invention itself, together with further objects andattendant advantages, will be best understood by reference to thefollowing detailed description, taken in conjunction with theaccompanying drawing.

Other features, benefits and advantages of the present invention willbecome apparent from the following description of the invention, whenviewed in accordance with the attached drawings and appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of this invention, reference shouldnow be made to embodiments illustrated in greater detail in theaccompanying figures and described below by way of examples of theinvention wherein:

FIG. 1 is a perspective view of an aircraft incorporating a sampleexterior fabric protective system in accordance with an embodiment ofthe present invention.

FIG. 2 is a side close-up cross-sectional view of the sample exteriorfabric protective system of FIG. 1.

FIG. 3 is a logic flow diagram illustrating a method of forming anexterior fabric protective structure in accordance with an embodiment ofthe present invention.

DETAILED DESCRIPTION

It has been determined through testing that under high use aircraftoperating conditions that lightning protective structures containingcarbon fiber with metal wires, which is disposed within an epoxy resintend to experience substrate microcracking and finish cracking. Thepresent invention overcomes this and is described in detail below. Whilethe present invention is described primarily with respect to theformation of an exterior protective structure for an aircraft, thepresent invention may be applied and adapted to various applications.The present invention may be applied in aeronautical applications, powerapplications, nautical applications, railway applications, automotivevehicle applications, medical applications, and commercial andresidential applications where the need for a durable lightningprotective structure that exhibits minimal or no weave telegraphing isdesired and particularly when weight or labor costs are of a concern.Also, a variety of other embodiments are contemplated having differentcombinations of the below described features of the present invention,having features other than those described herein, or even lacking oneor more of those features. As such, it is understood that the inventioncan be carried out in various other suitable modes.

In the following description, various operating parameters andcomponents are described for one constructed embodiment. These specificparameters and components are included as examples and are not meant tobe limiting.

Also, in the following description the term “component” refers to anartifact that is one of the individual parts of which a composite entityis made up. A component may refer to a part that can be separated fromor attached to a system, a part of a system or assembly, or other partknown in the art.

In addition, the term “surface” refers to the outer boundary of anartifact or a material layer constituting or resembling such a boundary.A surface may include not only the outer edge of a material, but also anoutermost portion or layer of a material. A surface may have a thicknessand include various particles.

Referring now to FIG. 1, a perspective view of an aircraft 10incorporating a sample exterior fabric protective system 12 inaccordance with an embodiment of the present invention is shown. Theprotective system 12 provides structural support across the exterior 14of the aircraft 10. The protective system 12 is applied over an aircraftpart(s), such as the fuselage 16 and tail or wings 18, of the aircraft10 to protect against lightning and to endure other environmentalconditions. The protective system in combination with the fuselage 16may be considered the main support structure of the aircraft 10. Theprotective system 12 includes multiple layers 20, which are described indetail below.

Referring now to FIG. 2, a side close-up cross-sectional view of theprotective system 12 in accordance with an embodiment of the presentinvention is shown. The protective system 12 includes a base substrate30. A hybrid prepreg fabric substrate 32 is disposed over and coupled tothe base substrate 30. A loaded surfacer film 52 is disposed over and iscoupled to the hybrid substrate 32. Finishes or finish layers 36, suchas spray-applied surfacers, pin hole fillers, primers, and painttopcoats, may be applied to the loaded surfacer 52. Although aparticular number of layers are shown for each of the base substrate 30,the hybrid substrate 32, and the loaded surfacer 52, any number oflayers may be utilized. In other words, although a single paint layer, asingle loaded surfacer layer, a single hybrid prepreg substrate layer,and four base substrate layers are shown, any number of each may beused. For example, additional layers of the hybrid substrate may be usedto improve performance.

The base substrate 30 may be a composite structure and include multiplestructural support layers 38. The structural support layers 38 may beformed of a carbon fiber/epoxy or the like and be in tape or fabricform. The structural support layers 38 may have varying angles ofplacement. For example, each adjacent layer, such as layers 40, 41 and42, may have unidirectional carbon tape in a different orientations,such as at 0.degree., +45.degree., and 90.degree., to increasestructural performance and overall durability of the base substrate 30.The base substrate 30 is the inner portion of the aircraft part. Theaircraft part may be a laminate, a sandwich structure, or a structurehaving a combination thereof and formed of various metallic ornonmetallic materials. Although in one envisioned embodiment of thepresent invention the base substrate 30 is applied to a carbon/epoxycomposite fuselage, the base substrate 30 may be applied to some othercomposite or hybrid parts formed of various materials known in the art.

The hybrid substrate 32 includes carbon fibers 44, metal integral to thecarbon in the form of wires 48 and an epoxy resin 46. FIG. 2 shows aresin rich area 47 around one wire 49, but actually the epoxy isdisposed throughout the hybrid substrate 32. The hybrid substrate 32 maybe in the form of an interwoven wire fabric (IWWF), as shown, or thelike. Fibers 44A are shown in a carbon tow of the hybrid substrate 32and extend perpendicular to the wires 48. Carbon tows 44B are shown andextend parallel to the wires 48. The fibers 44A and the carbon tows 44Btogether form an IWWF 53. Although wires 48 are shown as extending in asingle direction, they may extend in other directions.

The term “IWWF” refers to a carbon fiber that is woven typically into aplain weave fabric but other weave styles may be used. Each tow of theplain weave fabric contains a conductive component or wire, such as oneof the wires 48. In the embodiment shown, the wires 48 have anapproximate cross-sectional diameter D of 0.004 inches. The IWWF 53 doesnot include a discrete metal screen, as does metal woven fabrics orstructures having expanded metal foil layers. The IWWF 53 is a hybridfabric with metal wires contained therein as an integral component. TheIWWF 53 can carry load unlike the more common lightning protectionstructures containing metal meshes and foils. An example IWWF that maybe used is the Toray Composites (America) Inc. of Tacoma, Wash., IWWFepoxy prepreg FL6676G-37E containing intermediate modulus high strengthcarbon fibers. The IWWF 53 may have approximately 112 ksi tensionstrength and 10 MSI tension modulus when fabricated from IWWF prepregwith a 30%-50% epoxy resin content based on a carbon weight of about50%-70%. Another example that may be used to form the hybrid substrate32 is the combination of AS-4 carbon fiber from Hexcel Corporation ofDublin, Calif. with material identification no. 977 epoxy from CytecEngineered Materials Inc. of Anaheim, Calif.

The fibers within the hybrid substrate 32 may be of a variety of carbontypes. Of course, other fibers, fabric styles, metals and resins or thelike may be used that have similar properties to that of whichpreviously stated.

The hybrid substrate 32 can be pre-impregnated, more often referred toas “prepreg”, or can be a dry hybrid fabric product with epoxy resinadded as part of the production process such as with resin transfermolding or resin infusion. An example of a dry hybrid fabric product isAS-4 carbon fiber from Hexcel Corporation. A hybrid fabric substrate isprimarily carbon fabric with a small percentage of metal or other highlyconductive material coupled to the base substrate. The conductivecomponent of the hybrid substrate 32 discussed herein may be in the formof continuous metallic wire contained within approximately each carbontow 44B, but is not necessarily limited to that form or material. Theratio of metal/carbon/resin is dependent on the type of components andservice environment. The amount of resin utilized is maintained within apredetermined range to prevent microcracking and to maintain porositylevels below approximately 2% and to provide a desired structuralintegrity. The amount used may be measured with ultrasonic inspectionequipment. The microcracking and porosity levels set the low end of therange. The structural integrity desired sets the upper end of the range.

Unlike metal meshes, the hybrid substrate 32 provides structuralbenefit, not just electromagnetic protection. Weight of the hybridsubstrate that is associated with providing electromagnetic protectionincluding lightning protection, shielding protection, and staticelectricity protection is reduced because the amount of metal content isreduced.

The hybrid substrate 32 or the IWWF mentioned above may include wireformed of phosphor bronze, aluminum, nickel coated copper, copper,stainless steel, or other conductive materials having similar electricaland thermal characteristics or a combination thereof. Aluminum or othersimilar material, due to its density, conductivity, and thermalproperties may be used for improved lightning performance. On the otherhand, stainless steel or the like may be utilized for improved corrosionresistance. Cost, availability, corrosion susceptibility, internalstresses, including those from coefficient of thermal expansion, otherthermal and electrical properties are some of the parameters consideredwhen forming a protective structure for a given application. The hybridsubstrate 32 provides added structure support to the aircraft 10 and canbe used as part of the main support structure. The hybrid substrate 32may replace a portion of the base substrate 30 or structural supportlayers 38. Thus, the hybrid substrate 32, in general, reduces thethickness and weight of the base substrate 30, the time and costs toproduce the base substrate 30, and the overall weight of the aircraft10.

The loaded surfacer 52 is loaded with an inorganic filler (not shown),such as titanium dioxide. The inorganic fillers and a carrier,represented by the dashed line 51, are in the organic resin epoxy 50.The epoxy 50 may be loaded with inorganic fillers including fumed silicaand alumina, as well as other fillers known in the art or a combinationthereof. The carrier 51 may be in the form of a polyester mat, a carbonfiber mat, a glass mat, a metallized mat or the like. The amount ofconductive carriers may be increased to improve lightning resilience.The organic epoxy 50 is compatible to cure temperatures of approximatelybetween 250-350.degree. F. The loaded surfacer 52 has an approximateweight range between 0.02-0.06 pounds per square foot (lb/ft.sup.2) whenhaving an approximate thickness T of 0.004 inches above the wires 48after cure. The weight of the loaded surfacer 52 is proportionallyadjusted based on the resin content and flow characteristics of the IWWFor similar material and the tooling fit-up to ensure an adequate amountof surfacer is between the wire and outer surface for long termdurability in a high use commercial aircraft environment. An example ofa loaded surfacer that may be used is Cytec Engineered Materials SurfaceMaster 905, which has a nominal weight of 0.0325 lb/ft.sup.2. The loadedsurfacer 52 provides a surface that is suitable for priming andpainting.

Microscopic photographs similar to FIG. 2 reveal surfacer locallypenetrates through the hybrid substrate 32 down to the base substrate.The loaded surfacer 52 is selected such that the conductive component 44in the hybrid substrate 32 can vaporize through the loaded surfacer 52and any paint layers thereon for maximum protection.

Historically, surfacers have not been used because they tend to degradelightning performance of a protective structure. However, the commingledtechnique described herein along with the use of an IWWF provides aprotective structure that satisfies lightning protection requirements.The thickness T of the loaded surfacer 52 is adjusted depending upon theamount of lightning protection and the amount of other environmentalprotection desired. The amount of environmental protection, such as theamount of rain erosion resistance, the amount of corrosion resistance,and the overall durability, is traded against the amount of lightningprotection. In general, the thicker the loaded surfacer 52 the lesslightning protection, but the more environmental protection provided andvice versa. Continuing from the example above, the 0.0325 lb/ft.sup.2surfacer version using Surface Master 905 when combined with a hybridprepreg substrate with approximately 40% resin content, having a drypreform that consists of approximately 196 grams/meter.sup.2 of carbonand approximately 63 grams/meter.sup.2 of metal, can prevent puncturefrom initial lightning attachment for typical aircraft configurations. Atypical aircraft configuration is one with standard production finishesand that is tested to levels required for aircraft certification to meetFederal Aviation Regulation (FAR) Part 25.

The protective system 12 is durable and can withstand environmentalcycling associated with a commercial aircraft including those such ashigh use large commercial aircraft. Prior to approval for commercial useexterior portions of an aircraft undergo rigorous testing to simulatecommercial use. Some of this testing includes subjecting a component tolarge variations in temperature, humidity and pressure extremes. Forexample, a test specimen may be exposed to humidity levels of 95% forhours at approximately 120.degree. F. and may undergo 4000 or morecycles of temperature fluctuations from between approximately−65.degree. F. to 165.degree. F. to show commercial feasibility. Thetesting may include cycling for thousands of times in ground-air-groundchambers that simulate the change in pressures, humidity, andtemperature having limits that correspond with flight profiles.

Referring now to FIG. 3, a logic flow diagram illustrating a method offorming an exterior fabric protective structure or system, such as thesystem 12, in accordance with an embodiment of the present invention isshown. Although the following steps are primarily described with respectto the embodiment of FIG. 2, the steps may be easily modified to applyto other embodiments of the present invention.

In step 100, a loaded surfacer having a carrier, such as the loadedsurfacer 52 and the carrier 51, is applied to a mold. The mold may be ofvarious types, styles, shapes, and sizes, depending upon the componentor structure being formed. In step 102, a hybrid prepreg substrate, suchas the hybrid substrate 32 with carbon fabric 44 and metal 48, isapplied onto the loaded surfacer. In one embodiment, the prepregsubstrate is an IWWF, as described above. In step 104, a base substrate,such as the base substrate 30, is applied to the prepreg substrate. Inthis example described, the base substrate is in the form of prepreg andis co-cured. However, the layers of the base substrate may be preformedand cured prior to or subsequent to application of the prepregsubstrate. However, performance may be different depending on theco-bonding technique used.

The following steps 106-110 are similar to steps 100-104, however areperformed in a reverse order. In step 106, a base substrate is appliedto a mold in the form of a prepreg. The layers of the base substrate maybe preformed and cured prior to or subsequent to application on orinsertion into the mold. In step 108, a hybrid prepreg substrate havingcarbon fabric and an integral conductive component, such as wire, isapplied to the base substrate. In step 110, a loaded surfacer thatincludes a carrier is applied to the hybrid prepreg substrate.

In step 112, the loaded surfacer, the prepreg substrate, and the basesubstrate are cured to form the protective system. The mold includingthe loaded surfacer, the hybrid prepreg substrate, and the basesubstrate are placed into an autoclave or the like and baked. Thetemperature within the autoclave is approximately between245-355.degree. F. for this example. The curing temperature used isbelow the melting temperature of the conductive component or wireswithin the hybrid substrate to prevent melting of the wires. The lengthof time that the mold is within the autoclave is dependent upon thecuring properties of the materials utilized. During the curing processthe loaded surfacer is interfacially adhered to the hybrid prepregsubstrate rather than the loaded surfacer remaining entirely on top ofthe prepreg substrate. Matter or resin in the loaded surfacer and in theprepreg substrate mixes together and cures, thereby interfaciallyadhering to each other.

In addition, the loaded surfacer is such that when heated it evenlydistributes and cures across the mold and/or the prepreg substrate toform a single continuous layer having a certain thickness, such asthickness T. This assures that the wires, such as the wires 48, withinthe prepreg substrate are covered by the loaded surfacer. In an exampleembodiment, the loaded surfacer utilized cures to have a thickness thatis approximately 0.004 inches. The loaded surfacer may extend down tothe base substrate in localized areas. Commingling during cure of theresin in the surfacer with the resin in the hybrid fabric substratewhile maintaining sufficient surfacer layer content on top of theconductive component provides the desired environmental durability.

In step 114, upon curing of the protective structure or system formed bythe curing of the loaded surfacer, the prepreg substrate, and the basesubstrate is removed from the mold.

The above-stated commingled surfacer approach is especially beneficial,depending upon the application, for the higher strength and or highermodulus carbon types, particularly those that are impregnated withresins that are susceptible to microcracking. The above-described stepsare meant to be illustrative examples; the steps may be performedsequentially, synchronously, simultaneously, or in a different orderdepending upon the application. Of course, portions of the systemdescribed and the steps performed may be achieved manually or withoutthe use of specialized equipment or machines.

The present invention provides a cost effective and efficient system andmethod for the formation of lightning protective system. The presentinvention is lightweight, simplistic in design, prevents corrosion, andis durable. As such, the present invention increases service life andreduces maintenance costs of an aircraft and associated exteriorcomponents.

While the invention has been described in connection with one or moreembodiments, it is to be understood that the specific mechanisms andtechniques which have been described are merely illustrative of theprinciples of the invention, numerous modifications may be made to themethods and apparatus described without departing from the spirit andscope of the invention as defined by the appended claims.

1. An aircraft comprising an aircraft part and a protective fabricsystem for an exterior surface of the aircraft part, the protectivefabric system comprising: a base substrate; a hybrid prepreg interwovenwire fabric layer disposed over and coupled to the base substrate, thehybrid prepreg interwoven wire fabric layer comprising carbon fibers inthe form of tows woven into a weave pattern, an integral conductivecomponent in the form of metallic wire contained within approximatelyeach of said tows, and a first resin; a loaded surfacer disposed overand interfacially adhered to said hybrid prepreg interwoven wire fabriclayer, the loaded surfacer comprising a carrier loaded with at least oneinorganic filler and a second resin separate from the first resin,wherein in localized areas said loaded surfacer penetrates through saidhybrid prepreg interwoven wire fabric layer to said base substrate. 2.An aircraft as in claim 1 wherein said hybrid prepreg interwoven wirefabric layer provides structural support to the aircraft part.
 3. Anaircraft as in claim 1 wherein said loaded surfacer comprises inorganicfillers and forms a continuous layer over said hybrid prepreg interwovenwire fabric layer.
 4. An aircraft as in claim 3, wherein the continuouslayer has an approximate thickness of 0.004 inches.
 5. An aircraft as inclaim 1, further comprising matter adhered between said loaded surfacerand said hybrid prepreg interwoven wire fabric layer.
 6. An aircraft asin claim 1, wherein the aircraft part comprises a plurality ofstructural support layers that comprise the base substrate.
 7. Anaircraft as in claim 1 wherein said weave pattern comprises a plainweave and said first resin is an epoxy resin impregnated within saidweave pattern.
 8. An aircraft as in claim 1 wherein said hybrid prepreginterwoven wire fabric comprises a plurality of metallic wires disposedwith said tows.
 9. An aircraft as in claim 8 wherein said plurality ofmetallic wires are formed of at least one material selected fromphosphor bronze, nickel coated copper, copper, aluminum, and stainlesssteel.
 10. An aircraft as in claim 1 wherein said inorganic filler is amaterial selected from at least one of formed silica, alumina, andtitanium dioxide.
 11. An aircraft as in claim 1 wherein said secondresin comprises an epoxy having an approximate cure temperature between250-350° F.
 12. An aircraft as in claim 1 wherein said carrier is formedof material selected from the group consisting of a polyester mat, acarbon mat, a glass mat and a metallized mat.
 13. An aircraft as inclaim 1 wherein the base substrate is applied to the exterior surface ofthe aircraft part.
 14. An aircraft as in claim 1 wherein the first resinis disposed throughout the hybrid prepreg interwoven wire fabric layer.